1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to air cooled turbine blades.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine has a turbine section with a multiple stages of stationary vanes or nozzles and rotary blades or buckets exposed to extremely high temperature flow. The first stage vanes and blades are exposed to the highest temperature since the gas flow temperature progressively decreases through the turbine due to the extraction of energy. Especially in an industrial gas turbine engine, efficiency is the prime objective. In order to increase the efficiency of the engine, a higher gas flow temperature can be used in the turbine. However, the highest temperature that can be used depends upon the properties of the materials used in the turbine parts. For this reason, providing internal air cooling of the blades and vanes allows for a temperature higher than the material properties can withstand alone.
Another method of increasing the efficiency of the engine, for efficient use of the cooling air passing through the cooled airfoils is desired. Since the cooling air is generally bleed air from the compressor, maximizing the cooling effect while minimizing the amount of cooling air bled off from the compressor will increase the engine efficiency as well. Blade designers have proposed complex air cooling passages to maximize cooling efficiency while minimizing cooling volume. On a typical first stage turbine blade, the hottest surfaces occur at the airfoil leading edge, on the suction side immediately downstream from the leading edge, and on the pressure side of the airfoil at the trailing edge region. A showerhead arrangement is generally used to provide cooling for the leading edge of the airfoil. One problem blade designers are challenged with is that the hottest section on the suction side is also at a lower pressure than on the pressure side. A serpentine flow cooling circuit of the prior art that provides cooling for both the pressure side and the suction side will provide adequate cooling for the airfoil, but uses more cooling air that needed. Film cooling holes opening onto the pressure side and the suction side that are supplied with cooling air from the same cooling channel will both be discharging cooling air at the same pressure. Since the hot gas flow pressure on the suction side is lower than the pressure side, more cooling air will be discharged onto the suction side than is needed.
In a turbine airfoil with a serpentine flow cooling circuit, the cross sectional area of the passages must be sized in order than the airfoil walls will not be too thick. In many situations such as in open serpentine flow channels, some of the passages have cross sectional areas that are too large and result in low levels of heat transfer from the hot metal surface of the passage to the cooling air because the cooling air velocity is too low.
Turbine airfoils (which include blades and vanes) are typically cast as a single piece with the cooling passages cast within the airfoil. Ceramic cores having the cooling passage shape is used to form the airfoil.
It is an object of the present invention to provide a turbine airfoil with an internal cooling air circuit that would provide for a thermally balanced airfoil sectional temperature distribution.
It is another object of the present invention to provide for a turbine airfoil which allows for a maximize usage of the hot gas side pressure distribution in order to lower the required cooling air supply pressure to reduce the overall airfoil leakage flow.
It is another object of the present invention to provide for a ceramic core assembly with a minimum number of pieces while allowing for the above objectives to be met.